U.S. patent application number 14/684501 was filed with the patent office on 2016-10-13 for speed sensor for a gas turbine engine.
The applicant listed for this patent is United Technologies Corporation. Invention is credited to Jason Husband, William G. Sheridan.
Application Number | 20160298485 14/684501 |
Document ID | / |
Family ID | 57111287 |
Filed Date | 2016-10-13 |
United States Patent
Application |
20160298485 |
Kind Code |
A1 |
Sheridan; William G. ; et
al. |
October 13, 2016 |
SPEED SENSOR FOR A GAS TURBINE ENGINE
Abstract
A gas turbine engine includes a speed change mechanism. A fan
drive shaft has a radially extending surface that is attached to
the speed change mechanism. A speed sensor is located adjacent the
radially extending surface.
Inventors: |
Sheridan; William G.;
(Southington, CT) ; Husband; Jason; (South
Glastonbury, CT) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
United Technologies Corporation |
Hartford |
CT |
US |
|
|
Family ID: |
57111287 |
Appl. No.: |
14/684501 |
Filed: |
April 13, 2015 |
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F01D 21/003 20130101;
F01D 17/06 20130101; F05D 2260/40311 20130101; F02C 3/107
20130101 |
International
Class: |
F01D 21/00 20060101
F01D021/00; F02C 3/107 20060101 F02C003/107; F01D 17/06 20060101
F01D017/06 |
Claims
1. A gas turbine engine comprising: a speed change mechanism; a fan
drive shaft having a radially extending surface attached to the
speed change mechanism; and a speed sensor located adjacent the
radially extending surface.
2. The gas turbine engine of claim 1, wherein the radially
extending surface faces in an axial direction.
3. The gas turbine engine of claim 2, wherein the radially
extending surface is within 20 degrees of perpendicular to an axis
of rotation of the gas turbine engine.
4. The gas turbine engine of claim 1, wherein the radially
extending surface is located axially between a pair of fan drive
bearings.
5. The gas turbine engine of claim 1, wherein the speed sensor is
located axially upstream from the speed change mechanism.
6. The gas turbine engine of claim 1, wherein the speed sensor is
located adjacent a fan drive bearing.
7. The gas turbine engine of claim 1, wherein the speed sensor is
located axially between a pair of fan drive bearings.
8. The gas turbine engine of claim 1, further comprising a mount
flexibly supporting at least a portion of the speed change
mechanism.
9. A gas turbine engine comprising: a speed change mechanism; a fan
drive shaft having a radially extending surface attached to the
speed change mechanism; and a speed sensor located adjacent a fan
drive bearing.
10. The gas turbine engine of claim 9, wherein the speed sensor is
located axially between a pair of fan drive bearings.
11. The gas turbine engine of claim 9, wherein the speed sensor is
located adjacent a radially extending surface on the fan drive
shaft and the radially extending surface faces in an axial
direction.
12. The gas turbine engine of claim 11, wherein the radially
extending surface is within 20 degrees of perpendicular to an axis
of rotation of the gas turbine engine.
13. The gas turbine engine of claim 9, wherein the speed sensor is
located axially upstream from the speed change mechanism.
14. The gas turbine engine of claim 9, further comprising a mount
flexibly supporting at least a portion of the speed change
mechanism.
15. A method of designing a gas turbine engine comprising: locating
a speed sensor adjacent a fan drive shaft; positioning the speed
sensor relative to a fan drive shaft such that movement of the fan
drive shaft does not interfere with the operation of the speed
sensor.
16. The method of claim 15, further comprising attaching the fan
drive shaft to a speed change mechanism.
17. The method of claim 15, wherein the speed sensor is located
adjacent a radially extending portion of the fan drive shaft.
18. The method of claim 15, wherein the speed sensor is located
adjacent a fan drive bearing.
19. The method of claim 15, wherein the speed sensor is located
axially between a pair of fan drive bearings.
Description
BACKGROUND
[0001] One type of gas turbine engine includes a fan drive gear
system that is mechanically arranged between the turbo-machinery of
the engine and a fan. The turbo-machinery is composed of two
concentric shafts rotating at different speeds containing
independent compressors and turbines. The turbo-machinery
rotationally drives the fan, via the gear system, to move fluid
through a nacelle which divides the fluid flow into two streams. An
inner stream supplies the turbo-machinery and the outer stream
consists of fluid which bypasses the inner stream and is solely
compressed and moved by the fan.
[0002] Typically the fan drive gear system is provided by an
epicyclic gear train and includes a centrally located input gear
driven by the turbo-machinery, intermediate gears circumferentially
arranged about and intermeshing with the input gear and a ring gear
provided about and intermeshing the intermediate gears. Depending
upon the configuration, either the intermediate gears or the ring
gear rotationally drives the fan through a fan shaft in response to
rotation of the input gear.
[0003] The intermediate gears are typically supported in a carrier
by a journal extending between spaced apart walls of the carrier.
The carrier is typically constructed from a high strength metallic
alloy such as steel, titanium or nickel. The carrier is bolted to a
torque frame, which is secured to fixed structure or rotating
structure depending upon the particular type of gear system.
[0004] During operation of the gas turbine engine, a rotational
speed of the fan shaft is monitored with a speed sensor to ensure
proper functionality of the gas turbine engine. When the fan
experiences a load, such as during a bird strike, the fan shaft may
flex and change a distance between the speed sensor and the fan
shaft, which could damage the speed sensor. Therefore, there is a
need to reduce changes in the distance between the speed sensor and
the fan shaft.
SUMMARY
[0005] In one exemplary embodiment, a gas turbine engine includes a
speed change mechanism. A fan drive shaft has a radially extending
surface that is attached to the speed change mechanism. A speed
sensor is located adjacent the radially extending surface.
[0006] In a further embodiment of the above, the radially extending
surface faces in an axial direction.
[0007] In a further embodiment of any of the above, the radially
extending surface is within 20 degrees of perpendicular to an axis
of rotation of the gas turbine engine.
[0008] In a further embodiment of any of the above, the radially
extending surface is located axially between a pair of fan drive
bearings.
[0009] In a further embodiment of any of the above, the speed
sensor is located axially upstream from the speed change
mechanism.
[0010] In a further embodiment of any of the above, the speed
sensor is located adjacent a fan drive bearing.
[0011] In a further embodiment of any of the above, the speed
sensor is located axially between a pair of fan drive bearings.
[0012] In a further embodiment of any of the above, a mount
flexibly supports at least a portion of the speed change
mechanism.
In another exemplary embodiment, a gas turbine engine includes a
speed change mechanism. A fan drive shaft has a radially extending
surface that is attached to the speed change mechanism. A speed
sensor is located adjacent a fan drive bearing.
[0013] In a further embodiment of any of the above, the speed
sensor is located axially between a pair of fan drive bearings.
[0014] In a further embodiment of any of the above, the speed
sensor is located adjacent a radially extending surface on the fan
drive shaft. The radially extending surface faces in an axial
direction.
[0015] In a further embodiment of any of the above, the radially
extending surface is within 20 degrees of perpendicular to an axis
of rotation of the gas turbine engine.
[0016] In a further embodiment of any of the above, the speed
sensor is located axially upstream from the speed change
mechanism.
[0017] In a further embodiment of any of the above, a mount
flexibly supports at least a portion of the speed change
mechanism.
[0018] In another exemplary embodiment, a method of designing a gas
turbine engine includes locating a speed sensor adjacent a fan
drive shaft. The speed sensor is positioned relative to a fan drive
shaft such that movement of the fan drive shaft does not interfere
with the operation of the speed sensor.
[0019] In a further embodiment of any of the above, the fan drive
shaft is attached to a speed change mechanism.
[0020] In a further embodiment of any of the above, the speed
sensor is located adjacent a radially extending portion of the fan
drive shaft.
[0021] In a further embodiment of any of the above, the speed
sensor is located adjacent a fan drive bearing.
[0022] In a further embodiment of any of the above, the speed
sensor is located axially between a pair of fan drive bearings.
BRIEF DESCRIPTION OF THE DRAWINGS
[0023] FIG. 1 is a schematic view of an example gas turbine
engine.
[0024] FIG. 2 is a cross-sectional view of an example geared
architecture.
DETAILED DESCRIPTION
[0025] FIG. 1 schematically illustrates a gas turbine engine 20.
The gas turbine engine 20 is disclosed herein as a two-spool
turbofan that generally incorporates a fan section 22, a compressor
section 24, a combustor section 26 and a turbine section 28.
Alternative engines might include an augmentor section (not shown)
among other systems or features. The fan section 22 drives air
along a bypass flow path B in a bypass duct defined within a
nacelle 15, while the compressor section 24 drives air along a core
flow path C for compression and communication into the combustor
section 26 then expansion through the turbine section 28. Although
depicted as a two-spool turbofan gas turbine engine in the
disclosed non-limiting embodiment, it should be understood that the
concepts described herein are not limited to use with two-spool
turbofans as the teachings may be applied to other types of turbine
engines including three-spool architectures.
[0026] The exemplary engine 20 generally includes a low speed spool
30 and a high speed spool 32 mounted for rotation about an engine
central longitudinal axis A relative to an engine static structure
36 via several bearing systems 38. It should be understood that
various bearing systems 38 at various locations may alternatively
or additionally be provided, and the location of bearing systems 38
may be varied as appropriate to the application.
[0027] The low speed spool 30 generally includes an inner shaft 40
that interconnects a fan 42, a first (or low) pressure compressor
44 and a first (or low) pressure turbine 46. The inner shaft 40 is
connected to the fan 42 through a speed change mechanism, which in
exemplary gas turbine engine 20 is illustrated as a geared
architecture 48 to drive the fan 42 at a lower speed than the low
speed spool 30. The high speed spool 32 includes an outer shaft 50
that interconnects a second (or high) pressure compressor 52 and a
second (or high) pressure turbine 54. A combustor 56 is arranged in
exemplary gas turbine 20 between the high pressure compressor 52
and the high pressure turbine 54. A mid-turbine frame 57 of the
engine static structure 36 is arranged generally between the high
pressure turbine 54 and the low pressure turbine 46. The
mid-turbine frame 57 further supports bearing systems 38 in the
turbine section 28. The inner shaft 40 and the outer shaft 50 are
concentric and rotate via bearing systems 38 about the engine
central longitudinal axis A which is collinear with their
longitudinal axes.
[0028] The core airflow is compressed by the low pressure
compressor 44 then the high pressure compressor 52, mixed and
burned with fuel in the combustor 56, then expanded over the high
pressure turbine 54 and low pressure turbine 46. The mid-turbine
frame 57 includes airfoils 59 which are in the core airflow path C.
The turbines 46, 54 rotationally drive the respective low speed
spool 30 and high speed spool 32 in response to the expansion. It
will be appreciated that each of the positions of the fan section
22, compressor section 24, combustor section 26, turbine section
28, and fan drive gear system 48 may be varied. For example, gear
system 48 may be located aft of combustor section 26 or even aft of
turbine section 28, and fan section 22 may be positioned forward or
aft of the location of gear system 48.
[0029] The engine 20 in one example is a high-bypass geared
aircraft engine. In a further example, the engine 20 bypass ratio
is greater than about six (6), with an example embodiment being
greater than about ten (10), the geared architecture 48 is an
epicyclic gear train, such as a planetary gear system or other gear
system, with a gear reduction ratio of greater than about 2.3 and
the low pressure turbine 46 has a pressure ratio that is greater
than about five. In one disclosed embodiment, the engine 20 bypass
ratio is greater than about ten (10:1), the fan diameter is
significantly larger than that of the low pressure compressor 44,
and the low pressure turbine 46 has a pressure ratio that is
greater than about five 5:1. Low pressure turbine 46 pressure ratio
is pressure measured prior to inlet of low pressure turbine 46 as
related to the pressure at the outlet of the low pressure turbine
46 prior to an exhaust nozzle. The geared architecture 48 may be an
epicycle gear train, such as a planetary gear system or other gear
system, with a gear reduction ratio of greater than about 2.3:1. It
should be understood, however, that the above parameters are only
exemplary of one embodiment of a geared architecture engine and
that the present invention is applicable to other gas turbine
engines including direct drive turbofans.
[0030] A significant amount of thrust is provided by the bypass
flow B due to the high bypass ratio. The fan section 22 of the
engine 20 is designed for a particular flight condition--typically
cruise at about 0.8 Mach and about 35,000 feet. The flight
condition of 0.8 Mach and 35,000 ft (10,668 meters), with the
engine at its best fuel consumption--also known as "bucket cruise
Thrust Specific Fuel Consumption (`TSFC`)"--is the industry
standard parameter of 1 bm of fuel being burned divided by 1 bf of
thrust the engine produces at that minimum point. "Low fan pressure
ratio" is the pressure ratio across the fan blade alone, without a
Fan Exit Guide Vane ("FEGV") system. The low fan pressure ratio as
disclosed herein according to one non-limiting embodiment is less
than about 1.45. "Low corrected fan tip speed" is the actual fan
tip speed in ft/sec divided by an industry standard temperature
correction of [(Tram .degree.R)/(518.7 .degree.R)].sup.0.5. The
"Low corrected fan tip speed" as disclosed herein according to one
non-limiting embodiment is less than about 1150 ft/second (350.5
meters/second).
[0031] The example gas turbine engine includes fan 42 that
comprises in one non-limiting embodiment less than about twenty-six
(26) fan blades. In another non-limiting embodiment, fan section 22
includes less than about twenty (20) fan blades. Moreover, in one
disclosed embodiment low pressure turbine 46 includes no more than
about six (6) turbine rotors schematically indicated at 34. In
another non-limiting example embodiment low pressure turbine 46
includes about three (3) turbine rotors. A ratio between number of
fan blades 42 and the number of low pressure turbine rotors is
between about 3.3 and about 8.6. The example low pressure turbine
46 provides the driving power to rotate fan section 22 and
therefore the relationship between the number of turbine rotors 34
in low pressure turbine 46 and number of blades 42 in fan section
22 disclose an example gas turbine engine 20 with increased power
transfer efficiency.
[0032] As shown in FIG. 2, the geared architecture 48 is attached
to the inner shaft 40 through a flex shaft 60 to which an input
gear 62 (sun gear) is mounted for rotation about an axis A.
Intermediate gears 64 (in this example, star gears) are arranged
circumferentially about and intermesh with the input gear 62. A
ring gear 68 surrounds and intermeshes with the intermediate gears
64. Either the intermediate gears 64 or the ring gear 68
rotationally drives the fan drive shaft 66 depending upon the type
of epicyclic gear train configuration.
[0033] In the illustrated example, the geared architecture 48 is a
star gear system having the intermediate gears 64 rotationally
fixed relative to a rotational axis of the input gear 62. That is,
the intermediate gears 64 are permitted to rotate about their
respective rotational axes but do not rotate about the rotational
axis A of the input gear 62. The ring gear 68 is coupled to the fan
drive shaft 66 which rotationally drives the fan 42.
[0034] The engine static structure 36 of the gas turbine engine 20
includes a bearing compartment case 70 and a support member 72. A
torque frame 74 is affixed to the support member 72 which prevents
rotation of the torque frame 74 about the rotational axis A of the
input gear 62. Additionally, the support member 72 may be a
flexible support to aid in maintaining alignment of the geared
architecture 48 during operation of the gas turbine engine 20. In
the case where the geared architecture 48 is a planetary gear
configuration, the torque frame 74 would rotate about the
rotational axis A and the ring gear 68 would be coupled to the
engine static structure 36 with a rigid or flexible support.
[0035] The torque frame 74 includes multiple shafts 76 integral
with a base 78 that provide first and second support features 80,
82 affixed to the support member 72. Each shaft 76 includes a
bearing assembly 84 for rotationally supporting its respective
intermediate gear 64. In one example, the torque frame 74 includes
five equally circumferentially spaced shafts 76 that
correspondingly support five star or intermediate gears 64. The
base 78 and shafts 76 of the torque frame 74 are unitary and formed
by a one-piece structure, for example, by a cast steel structure.
Other high strength metallic alloys, such titanium or nickel, may
also be used.
[0036] During operation of the gas turbine engine 20, a speed
sensor 90 is used to monitor a rotational speed of the fan drive
shaft 66. The speed sensor 90 is used to verify that the fan drive
shaft 66 is rotating at an appropriate rotational speed relative to
the inner shaft 40. A significant deviation from the predicted
relative rotational speeds between the fan drive shaft 66 and the
inner shaft 40 can indicate an uncoupling between the fan drive
shaft 66 and the inner shaft 40 or a mechanical issue with the
geared architecture 48. Therefore, having an accurate reading of
the rotational speed of the fan drive shaft 66 is important when
monitoring the health of the gas turbine engine 20.
[0037] In order for the speed sensor 90 to function properly and
accurately measure the rotational speed of the fan drive shaft 66,
a very small gap between the speed sensor 90 and the fan drive
shaft 66 must be maintained. In the illustrated example, the gap
extends a distance D1 between a distal end of the speed sensor 90
and a portion of the fan drive shaft 66. In one example, the speed
sensor 90 is a magnetic pickup sensor and the distance D1 is
approximately 0.030 inches (approximately 0.762 mm) However,
additional types of speeds sensors 90 could be used in place of the
magnetic pickup sensor.
[0038] During operation of the gas turbine engine 20, the geared
architecture 48 is allowed to move on flexible supports to maintain
alignment of the components in the geared architecture 48. The
flexible supports include the flex shaft 60, a flexible ring gear
support 69, the support member 72. In one example, the geared
architecture 48 is allowed to move approximately 0.050 inches (1.27
mm) in a radial direction on the flexible supports.
[0039] When the gas turbine engine 20 experiences high loads, such
as during bird strikes, the geared architecture 48 may flex as
discussed above. During a bird strike, the geared architecture 48
can flex and move out of alignment with the speed sensor 90 or
directly contact and possibly damage the speed sensor 90.
[0040] In one example, the speed sensor 90 is located adjacent a
radially extending portion 66A of the fan drive shaft 66. The
radially extending portion 66A includes a radially extending
surface 67 that faces an axial direction. The radially extending
surface 67 is within 20 degrees of perpendicular to the axis A of
rotation of the gas turbine engine 20. By locating the speed sensor
90 adjacent the radially extending portion 66A of the fan drive
shaft 66, the geared architecture 48 is able to flex under a load,
and the radially extending portion 66A of the fan drive shaft 66
moves in a generally radial direction parallel to a distal end of
the speed sensor 90 and avoids contact with the speed sensor
90.
[0041] The speed sensor 90 may also be located adjacent a fan drive
bearing 92. The closer the speed sensor 90 is to the fan drive
bearing 92 the less movement the fan drive shaft 66 will have
relative to the speed sensor. Although the speed sensor 90 is
located axially forward of the geared architecture 48 in the
illustrated example, the speed sensor 90 could also be located
downstream of the geared architecture 48 if the low pressure
compressor 44 is tied to the fan 42.
[0042] In another example, a speed sensor 90A is located axially
between a pair of the fan drive bearings 92. The speed sensor 90A
is similar to the speed sensor 90 except where described below or
shown in the Figures. The speed sensor 90A may be located closer to
one of the two fan drive bearings 92 or directly between the two
fan drive bearings 92. A distal end of the speed sensor 90A is
located adjacent a radially extending flange 96 on an axially
extending portion 66B of the fan drive shaft 66. The radially
extending flange 96 includes a radially extending surface 98 that
faces in an axial direction. The radially extending surface 98 is
within is within 20 degrees of perpendicular to the axis A of
rotation of the gas turbine engine 20.
[0043] In yet another example, a speed sensor 90B is also located
axially between the pair of fan drive bearings 92. The speed sensor
90B is similar to the speed sensor 90 except where described below
or shown in the Figures. The speed sensor 90B may be located closer
to one of the two fan drive bearings 92 or directly between the two
fan drive bearings 92. A distal end of the speed sensor 90B is
located adjacent the axially extending portion 66B of the fan drive
shaft 66.
[0044] Although the three speed sensors 90, 90A, and 90B are shown
in the illustrated example, a single one of the speed sensors 90,
90A, and 90B could be utilized or any combination of speed sensors
90, 90A, and 90B could be utilized.
[0045] The preceding description is exemplary rather than limiting
in nature. Variations and modifications to the disclosed examples
may become apparent to those skilled in the art that do not
necessarily depart from the essence of this disclosure. The scope
of legal protection given to this disclosure can only be determined
by studying the following claims.
* * * * *